CN114837807A

CN114837807A - Aircraft propulsion system with inter-turbine combustor

Info

Publication number: CN114837807A
Application number: CN202210101168.6A
Authority: CN
Inventors: 杰弗里·道格拉斯·兰博
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2021-02-01
Filing date: 2022-01-27
Publication date: 2022-08-02
Also published as: US11859539B2; US20220243667A1

Abstract

An aircraft propulsion system and computing system are provided. The propulsion system includes a Low Pressure (LP) spool and a core engine having a High Pressure (HP) spool. The frame is positioned between the HP turbine and the LP turbine in a serial flow arrangement. The frame includes an inter-turbine combustor that includes struts that form exit openings into a core flow path of the propulsion system. The first fuel system is configured to flow a liquid fuel to the combustion section to produce a first combustion gas. The second fuel system is configured to flow a gaseous fuel to the core flow path via the inter-turbine combustor to generate a second combustion gas. The propulsion system forms a rated power output ratio of the core engine to the inter-turbine combustor with the LP spool between 1.5 and 5.7.

Description

Aircraft propulsion system with inter-turbine combustor

Technical Field

The present subject matter relates generally to aircraft propulsion systems. The present subject matter is particularly concerned with structures and methods for engine operation of aircraft propulsion systems.

Background

Conventional aircraft propulsion systems are typically configured to generate all levels of thrust from combustion gases from a combustion system positioned between a High Pressure Compressor (HPC) and a High Pressure Turbine (HPT). Thus, the HPC, combustion system, and HPT are sized to produce the full range of thrust outputs or maximum thrust outputs.

Some propulsion systems include a reheat system (e.g., an intensifier or afterburner) to generate an increased amount of thrust. However, such systems are often inefficient in terms of fuel consumption, and further generate emissions or noise that exceed the regulated levels of emissions and noise, such as commercial and general aviation aircraft. Reheat systems for non-aircraft gas turbine engines, such as industrial gas turbines used for power generation, do not require consideration of propulsion efficiency and overall aircraft weight, performance, and efficiency.

Accordingly, there is a need for an improved aircraft propulsion system that can generate a large amount of thrust without adversely affecting emissions output and fuel consumption.

Disclosure of Invention

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

Aspects of the present disclosure relate to an aircraft propulsion system. The propulsion system includes a Low Pressure (LP) spool including a fan section, an LP compressor, and an LP turbine. The core engine includes a High Pressure (HP) compressor, a combustion section, and an HP turbine. The HP compressor and HP turbine together form a rotatable HP spool. The frame is positioned between the HP turbine and the LP turbine in a serial flow arrangement. The frame includes an inter-turbine combustor that includes struts that form exit openings into a core flow path of the propulsion system. The first fuel system includes a first fuel conduit in fluid communication with the fuel nozzle at the combustion section and configured to flow a liquid fuel to the combustion section to produce a first combustion gas. The second fuel system includes a second fuel conduit in fluid communication with the core flow path at the inter-turbine combustor via an outlet opening and configured to flow a gaseous fuel to the core flow path to produce a second combustion gas. The LP compressor, HP compressor, combustion section, HP turbine, inter-turbine combustor, and LP turbine are in a serial flow arrangement. The propulsion system forms a rated power output ratio of the core engine to the inter-turbine combustor with the LP spool between 1.5 and 5.7.

Another aspect of the present disclosure relates to a computing system for an aircraft propulsion system. The computing system includes one or more processors and one or more memories, where the memories are configured to store instructions that, when executed by the processors, cause the propulsion system to operate. The operation includes: flowing a liquid fuel stream to a combustion section of the propulsion system; producing a first combustion gas in the combustion section corresponding to 85% or less of a rated power output of the propulsion system; and modulating a rotational speed of the LP spool via modulating a flow of gaseous fuel to the inter-turbine combustor to produce a second combustion gas.

These and other features, aspects, and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description, serve to explain the principles of the invention.

Drawings

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is an exemplary embodiment of an aircraft including a propulsion system according to aspects of the present disclosure;

FIG. 2 is a schematic cross-sectional view of a propulsion system for the aircraft of FIG. 1, according to aspects of the present disclosure;

FIG. 3 is a schematic cross-sectional view of a portion of a propulsion system including an embodiment of an inter-turbine combustor in accordance with aspects of the present disclosure; and

FIG. 4 is a flowchart outlining the steps of a method for operating a propulsion system according to aspects of the present disclosure.

Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the invention.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one element from another, and are not intended to denote the position or importance of the various elements.

The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

Embodiments of an aircraft and a propulsion system including a turbine-to-turbine reheat combustor positioned between a first turbine and a second turbine are provided. The first fuel system provides liquid fuel to the combustion section to produce combustion gases to the turbine. The second fuel system provides gaseous fuel to the inter-turbine combustor to selectively produce the reheat gas based on specific engine or aircraft operating conditions. Embodiments of the propulsion system are also configured to produce a particular thrust or power output ratio based on the first fuel system in conjunction with the first fuel system and the second fuel system.

Embodiments of the propulsion system and aircraft provided herein allow for relatively small core engine sizes (i.e., the size and power output of the high-pressure spool, combustion section, and high-pressure turbine together), while producing a rated power output similar to the larger core engine size via increasing the power extracted from the low-pressure spool. Smaller core engine sizes allow for reduced fuel consumption, reduced emissions, greater bypass ratios, and improved specific fuel consumption rates. The smaller core engine size also allows the propulsion system to be operated as an Auxiliary Power Unit (APU) to power aircraft subsystems, electronics, or to provide engine starting power for other propulsion systems without the use of a dedicated APU separate from the propulsion system. Such a system allows for improved overall aircraft efficiency, for example, by eliminating the need or desire for a non-propulsive gas turbine engine.

Referring now to the drawings, in fig. 1, an exemplary embodiment of a vehicle 100 is provided, the vehicle 100 including a propulsion system 10 having an inter-turbine combustor in accordance with aspects of the present disclosure. In an embodiment, the vehicle 100 is an aircraft that includes an aircraft structure or airframe (airframe) 105. The body 105 includes a fuselage 110, to which are attached wings 120 and an empennage 130. A propulsion system 10 according to aspects of the present disclosure is attached to one or more portions of the airframe. In various embodiments, aircraft 100 includes a thermal management system 200 configured to desirably distribute thermal loads, such as adding or removing heat from one or more fluids or structures (such as, but not limited to, oxidants at propulsion systems, fuels for electric machines, electronics, computing systems, environmental control systems, gear assemblies, or other systems or structures, lubricants, hydraulic fluids, pneumatic fluids, or cooling fluids).

In various embodiments, the aircraft 100 includes subsystems that generally define electrical loads that require input energy. Such a system includes an anti-icing system 160, an environmental control system 150, and an avionics system 140. The propulsion system 10 is configured to extract energy from one or more spools to power aircraft subsystems, such as described herein. Although some systems may be formed as mechanical systems, electrification of the systems may reduce the weight and complexity of the aircraft. However, such electrification typically requires greater energy input, such as from the propulsion system 10 described herein.

In some cases, propulsion system 10 is attached to the rear of fuselage 110. In certain other instances, propulsion system 10 is attached below, above, or through a portion of wing 120 and/or empennage 130. In various embodiments, propulsion system 10 is attached to airframe 105 via a pylon or other mounting structure. In other embodiments, the propulsion system 10 is housed within an airframe, such as may be illustrated in certain supersonic commercial aircraft.

Referring now to FIG. 2, a schematic cross-sectional view of a propulsion system for an aircraft according to an exemplary embodiment of the present disclosure is provided. As shown in FIG. 2, propulsion system 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about axial direction A; not depicted). In various embodiments, propulsion system 10 is configured as a gas turbine engine, such as a turbofan engine. In particular embodiments, propulsion system 10 is a ductless open rotor engine (i.e., without a nacelle surrounding fan blades). Generally, the propulsion system 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

The exemplary turbine 16 illustrated generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The housing 18 encloses in serial flow relationship: a compressor section including a first booster or Low Pressure (LP) compressor 22 and a second High Pressure (HP) compressor 24; a combustion section 26; a turbine section including a first High Pressure (HP) turbine 28 and a second Low Pressure (LP) turbine 30; and an injection exhaust nozzle section 32. A High Pressure (HP) shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. A Low Pressure (LP) shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and injection exhaust nozzle section 32 are arranged in a serial flow sequence and together define a core air flow path 37 through the turbine 16.

In certain embodiments, propulsion system 10 includes one or more electric motors 370 that are operably coupled to the spool of the engine. The motor 370 may be operably coupled to the HP spool, the LP spool, or both to extract or receive energy from the spools during operation. Further, the motor 370 may be configured to output or release energy to the spool to initiate or assist in the rotation of the HP spool (e.g., during start-up or other desired operations), or to output or release energy to the LP spool during desired operations of the aircraft (e.g., during cruise operations, or transient conditions, or relative bursts of thrust or power output). In various embodiments described herein, the HP spool may be allowed to operate in a substantially steady state condition, for example, to allow substantially steady state extraction of energy to the motor. The motors may release energy to one or more subsystems (e.g.,

subsystems

140, 150, 160) at the aircraft 100. In particular, embodiments of the propulsion system 10 provided herein allow for increased energy extraction from the HP spool. Still further or alternatively, the system 10 may allow for power extraction during ground operating conditions (including ground idle or coasting conditions).

In particular embodiments, such as depicted in FIG. 2, fan section 14 may include a variable pitch fan 38. The turbine 16 is operatively coupled to the fan 38 to drive the fan 38. The fan 38 includes a plurality of rotatable fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, fan blades 40 extend generally outward from disk 42 in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to a suitable actuating member 44, which actuating member 44 is configured to collectively change the pitch of the fan blades 40, e.g., in unison. Fan blades 40, disks 42, and actuating members 44 rotate together about longitudinal axis 12 via LP shaft 36 across power gearbox 46. Power gearbox 46 includes a plurality of gears for reducing the speed of LP shaft 36 to a more efficient fan speed. Thus, for the depicted embodiment, the turbine 16 is operatively coupled to the fan 38 through the power gearbox 46.

Still referring to FIG. 2, the compressed second portion of air 64 from the compressor section is mixed with a liquid fuel and combusted within the combustion section to provide combustion gases 66. Combustion gases 66 are routed (routed) from combustion section 26 through HP turbine 28, wherein a portion of the thermal and/or kinetic energy from combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 coupled to casing 18 and HP turbine rotor blades 70 coupled to HP shaft 34, thereby rotating HP shaft 34, thereby supporting operation of HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30, wherein a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thus rotating the LP shaft 36, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38.

The combustion gases 66 are then routed through the injection exhaust nozzle section 32 of the turbine 16. At the same time, as first portion of air 62 is routed through bypass airflow passage 56 before being discharged from fan nozzle exhaust section 76 of propulsion system 10, the pressure of first portion of air 62 increases significantly, also providing propulsion thrust. HP turbine 28, LP turbine 30, and jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing combustion gases 66 through turbine 16.

It should be appreciated that the exemplary propulsion system 10 depicted in FIG. 2 is a relatively high power class turbofan propulsion system 10. Thus, when operating at a rated speed, propulsion system 10 may be configured to generate a relatively large amount of thrust. More particularly, when operating at a rated speed, propulsion system 10 may be configured to produce at least about 14,000 pounds of thrust, or at least 18,000 pounds of thrust, or at least 21,000 pounds of thrust, or at least 24,000 pounds of thrust, or at least 30,000 pounds of thrust. Some embodiments may produce up to 120,000 pounds of thrust at rated speeds. Thus, the propulsion system 10 depicted in FIG. 2 may be referred to as a medium to high power class gas turbine engine.

It should be appreciated that other exemplary embodiments of the propulsion system 10 are relatively high power class turboshaft propulsion systems 10. Thus, the propulsion system 10 may be configured to produce a relatively large amount of horsepower when operating at rated speeds. More specifically, propulsion system 10 may be configured to produce up to 10,000 shaft horsepower (shp) when operating at rated speed. In various embodiments, propulsion system 10 may be configured to produce at least 2,000 shaft horsepower (shp) when operating at rated speed.

Further, it should be appreciated that the exemplary propulsion system 10 depicted in FIG. 2 is by way of example only, and that in other exemplary embodiments, the propulsion system 10 may have any other suitable configuration. For example, in certain exemplary embodiments, the fan may not be a variable pitch fan. Additionally or alternatively, aspects of the present disclosure may be used with any other suitable aircraft gas turbine engine (e.g., turboshaft engine, turboprop engine, turbojet, etc.). Further embodiments may omit the nacelle surrounding the fan blades, for example to form an open rotor turbofan engine.

It should be understood that, as used herein, rotation and modulation of the speed of the HP spool and the LP spool corresponds to generation and modulation of output torque, power, or thrust. In a turbofan configuration of a propulsion system, most of the thrust is generated by the rotation of the fan blades via the LP spool. In various embodiments, the remaining portion of thrust is generated via combustion gases discharged through an exhaust gas injection nozzle.

Referring now to fig. 3, a close-up view of a portion of the exemplary propulsion system 10 of fig. 2 is provided. More specifically, FIG. 3 provides a close-up view of the combustion section 26 and the turbine section. In a particular embodiment, the combustion section 26 includes a combustor assembly 100. The combustor assembly 100 may be configured as a detonation combustor assembly, such as, but not limited to, an annular combustor, a dual annular combustor, a can combustor, a trapped vortex combustor, or other suitable combustion system. The burner assembly may be configured as a lean-burn burner, a rich-lean-quench (RQL) burner, or other suitable burner assembly.

In one embodiment, the combustion section 26 includes a first fuel conduit, for example, formed by one or more fuel nozzles 124, the one or more fuel nozzles 124 configured to receive a flow of liquid fuel (as schematically shown by arrow 352) and provide the liquid fuel to the combustion chamber 114 for combustion or detonation. Although not depicted in further detail, the fuel nozzle 124 may be any suitable type of fuel injector, nozzle, rail, or other liquid fuel dispensing, atomizing, or mixing device. In particular embodiments, the fuel nozzles 124 may be configured for lean or rich mixtures, combustion, or knock.

In certain embodiments, combustor assembly 100 generally includes an inner liner 102 extending generally along axial direction a between an aft end and a forward end, and an outer liner 108 also extending generally along axial direction a between the aft end and the forward end. Together, the inner liner 102 and the outer liner 108 at least partially define a combustion chamber 114 therebetween. The inner liner 102 and the outer liner 108 are each attached to or integrally formed with the annular dome. More specifically, the annular dome includes an inner dome section 116 integrally formed with the front end 106 of the inner liner 102 and an outer dome section 118 generally formed with the front end of the outer liner 108. Further, the inner and

outer dome sections

116, 118 may each be integrally formed (or alternatively may be formed from multiple components attached in any suitable manner), and may each extend in the circumferential direction C to define an annular shape.

However, it should be understood that in other embodiments, combustor assembly 100 may not include inner dome section 116 and/or outer dome section 118; may include separately formed inner dome section 116 and/or outer dome section 118 attached to respective inner and

outer liners

102, 108; or may have any other suitable configuration. In other embodiments, combustion section 26 may be configured as a detonation combustion system, such as a rotary detonation combustion system or a pulse detonation combustion system.

Still referring to fig. 3, the combustor assembly 100 also includes a plurality of fuel-air mixers spaced apart in a circumferential direction C (not shown) and positioned at least partially within the annular dome. More specifically, a plurality of fuel-air mixers is at least partially disposed between the outer dome section 118 and the inner dome section 116 in the radial direction R. Compressed air from the compressor section of propulsion system 10 flows into or through a fuel-air mixer where it is mixed with fuel and ignited to generate combustion gases 66 within combustion chamber 114. Inner dome section 116 and outer dome section 118 are configured to help provide such a flow of compressed air from the compressor section into or through fuel-air mixer 124. For example, outer dome section 118 may include an outer shroud at the forward end, while inner dome section 116 similarly includes an inner shroud at the forward end. The outer and inner shrouds may help direct the flow of compressed air from the compressor section into or through one or more fuel-air mixers. However, again in other embodiments, the annular dome may be configured in any other suitable manner.

Certain embodiments of the combustion section 26 or turbine section may include one or more components formed from Ceramic Matrix Composite (CMC) materials. In certain embodiments, both the inner liner 102 and the outer liner 108 are formed of a CMC material. In certain embodiments, the vanes or struts of the frame 300, described further below, are formed from a CMC material. Still further embodiments include one or more bucket or blade stages of the LP turbine 30 formed from the CMC material. CMC materials are non-metallic materials with high temperature capabilities. Exemplary CMC materials for such components may include silicon carbide (SiC), silicon nitride, or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded in a matrix, such as oxidation-stable reinforcing fibers, including monofilaments such as sapphire and silicon carbide (e.g., SCS-6 of Deslon), and including silicon carbide (e.g., of Nippon Carbon)

Of Ube Industries

And Dow Corning

) Aluminum silicates (e.g., 440 and 480 of Nextel) and chopped whiskers and fibers (e.g., 440 and 480 of Nextel)

) And optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in some embodiments, theThe fiber bundles including the ceramic refractory coating are formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple ribbons can be stacked together (e.g., as plies) to form a preform component. The fiber bundles may be impregnated with the slurry composition prior to forming the preform or after forming the preform. The preform may then be heat treated (e.g., cured or burned) to produce a high char residue in the preform, and subsequently chemically treated (e.g., melt infiltrated with silicon) to obtain a component formed from the CMC material having the desired chemical composition. In other embodiments, the CMC material may be formed as, for example, a carbon fiber cloth instead of a tape. Additionally or alternatively, the CMC material may be formed in any other suitable manner or using any other suitable material.

Still referring to FIG. 3, and as discussed above and further below, combustion gases 66 flow from combustors 114 into and through the turbine section of propulsion system 10, wherein a portion of the thermal and/or kinetic energy from combustion gases 66 is extracted via sequential stages of turbine stator vanes and turbine rotor blades within HP turbine 28 and LP turbine 30. More specifically, as shown in FIG. 3, the combustion gases 66 from the combustor 114 flow into the HP turbine 28 located immediately downstream of the combustor 114, wherein thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 and HP turbine rotor blades 70.

As also discussed above with reference to FIG. 2, the HP turbine 28 is coupled to the HP compressor 24 via the HP shaft 34 to form a HP spool or HP spool that is operable to a maximum speed that is generally higher than the LP spool formed by the LP compressor 22, LP turbine 30, LP shaft 36, and fan section 14. Thus, rotation of the multi-stage HP turbine rotor blades 70 correspondingly rotates the multi-stage HP compressor rotor blades 80.

2-3 are configured to operate to maintain the temperature of HP turbine 28 below the maximum operating temperature of the various components therein, while reducing the cooling flow extracted from the compressor section. In a particular embodiment, the HP turbine includes one or more stages of blades formed as substantially solid, impermeable blades at the airfoil at the core flow path. In other embodiments, the HP turbine includes one or more stages of blades configured to reduce cooling flow therethrough via reducing or eliminating cooling flow from the compressor section to the HP turbine, such as to improve engine efficiency by reducing or eliminating the amount of air removed from the thermodynamic cycle (i.e., air removed from combustion).

Referring back to FIG. 3, the turbine section includes an inter-turbine frame 300 positioned between the HP turbine 28 and the LP turbine 30. Frame 300 is configured as a stationary, static support structure configured to support one or both of HP turbine 28 or LP turbine 30. Frame 300 includes an inter-turbine combustor 310 configured to allow a flow of gaseous fuel into a core flow path upstream of the LP turbine. The frame forms a combustor at the vanes or struts 312 of the frame and one or more apertures or openings 306, the one or more apertures or openings 306 configured to allow a flow of gaseous fuel 362 through the struts 312 into the core flow path.

In various embodiments, the inter-turbine combustor 310 forms a second fuel conduit configured to deliver gaseous fuel 362 to the core flow path. In particular embodiments, in contrast to combustion systems formed at combustion sections, for example, combustors 310 are formed as flame holders at struts 312 of frame 300. In certain embodiments, the inter-turbine combustor includes struts or buckets formed, for example, to generally provide airfoils and/or structural members for an inter-turbine frame, mid-frame structure, or other support frame. The strut 312 includes a forward or leading edge 304 and an aft or trailing edge 302. The struts 312 include hollow portions to allow fluid flow therethrough. In certain embodiments, frame 300 includes a lubricant conduit 316 and an air conduit 314, for example, that generally provide lubricant or air for bearing assembly 320. Inter-turbine frame 300 may also include a duct 308 configured to exhaust the flow of gaseous fuel 362 through apertures 306 at struts 312. In a particular embodiment, apertures 306 are positioned at trailing edge 302 of struts 312 to allow gaseous fuel to flow therethrough and aft toward LP turbine 30.

The combustion section 26 is configured as a deflagration or detonation combustion section. Liquid fuel352 is provided to the combustion section 26 by one or more fuel nozzles 124. The flow of liquid fuel 352 is mixed with compressed air from the compressor section and then combusted to produce combustion gases 66. The liquid fuel 352 provided to the combustion section 26 is a liquid Jet fuel or an aviation turbine fuel, such as a kerosene-based fuel, a naphtha-based fuel, or an equivalent (e.g., Jet-A, Jet-B, JP8, a biofuel, a synthetic fuel, or other suitable aviation fuel). The gaseous fuel 362 provided to the inter-turbine combustor is a gaseous fuel, such as hydrogen (H) ₂ ) Natural gas, methane, synthesis gas, or other suitable types of gaseous fuels. The flow of gaseous fuel 362 released through inter-turbine combustor 310 between HP turbine 28 and LP turbine 30 mixes with the flow of combustion gases 66.

It should be appreciated that gaseous fuel 362 has a gaseous fuel ignition temperature (i.e., a second ignition temperature) that is less than the liquid fuel ignition temperature (i.e., a first ignition temperature) of liquid fuel 352. Gaseous fuel 362 also has a gaseous fuel combustion rate (i.e., a second combustion rate) that is greater than the liquid fuel combustion rate (i.e., a first combustion rate) of liquid fuel 352. The relatively lower second firing temperature limit allows for a mixture of gaseous fuel 362 from the inter-turbine combustor 310 and combustion gases 66 from the combustion section 26 to generate the second combustion gases with a relatively high velocity fluid flow through the turbine section. The flame speed of the second combustion gas produced by the inter-turbine combustor 310 may also be greater than the flame speed of the first combustion gas produced by the combustion section 26.

In certain embodiments, the upper flammability limit of the gaseous fuel 362 is greater than the liquid fuel. Further, in certain embodiments, the range of flammability limits is generally greater than the range of liquid fuels. In certain embodiments, the lower flammability limit of the gaseous fuel is less than the upper flammability limit of the liquid fuel. Still further, gaseous fuels have a higher degree or magnitude of flammability than liquid fuels. Thus, unlike afterburner systems that utilize liquid fuels, a mixture of gaseous fuel and combustion gases can be combusted without external ignition (e.g., using an igniter or other energy input).

The aircraft 100 and the propulsion system 10, individually or together, include a first fuel system 350 for flowing and distributing liquid fuel 352 at the combustion section 26, and a second fuel system 360 for flowing and distributing gaseous fuel 362 at the inter-turbine combustor 310. It should be appreciated that first fuel system 350 may be further configured to provide liquid fuel 352 as an actuating fluid and/or a heat exchange fluid (e.g., to receive heat or thermal energy from another fluid or surface) as opposed to second fuel system 360. More specifically, the first fuel system 350 may be configured to provide an actuation force or pressure to modulate one or more valves, actuators, doors, openings, nozzles, flow devices, or adjustable areas at the propulsion system, such as variable area nozzles, bleed valves, exhaust nozzles, active clearance control valves or doors, transient or start bleed valves, or other actuatable portions of the propulsion system or aircraft.

The embodiments of the aircraft 100 and propulsion system 10 depicted and described herein may provide improved propulsion system and aircraft efficiency, emissions, or fuel burn. The inter-turbine combustor 310 may increase the LP turbine 30 power extraction for a given High Pressure (HP) spool or core engine size (i.e., HP compressor 24, combustion section 26, and HP turbine 38). The separation of the second fuel system 360 configured to provide gaseous fuel 362 to the inter-turbine combustor 310 from the first fuel system 350 configured to provide liquid fuel 352 to the combustion section 26 allows for increased LP turbine power extraction and power output that is greater than the output power from the core engine alone.

It should be appreciated that although described as an inter-turbine combustor between a HP turbine and a LP turbine, various embodiments provided herein may include an inter-turbine combustor between a first turbine that receives higher pressure combustion gases and a second turbine that receives lower pressure combustion gases. Accordingly, various embodiments may include an Intermediate Pressure (IP) turbine generally positioned between the HP turbine and the LP turbine. Particular embodiments may position the inter-turbine combustor described herein between the HP turbine and the IP turbine, or between the IP turbine and the LP turbine.

Still further, although described as a conventional turbine rotor, embodiments of the HP or LP turbines provided herein may be configured as interdigitated or vaneless turbine assemblies.

Embodiments of aircraft 100 and propulsion system 10 provided herein allow for sizing and operating core engines at steady state speeds and power output, particularly for hybrid electric propulsion systems, and/or avoid power generation from a separate Auxiliary Power Unit (APU). In certain embodiments, propulsion system 10 is configured to generate a work distribution between a core engine including a High Pressure (HP) spool and a combustion section and a Low Pressure (LP) spool including an inter-turbine combustor. In various embodiments, propulsion system 10 has a rated power output ratio of between 1.5 and 5.7 for the core engine to the inter-turbine combustor 310 with the LP spools (i.e., LP turbine 30, LP compressor 22, and fan section 14). In certain embodiments, the propulsion system is configured to produce 85/15 work distribution between the core engine and the LP spool. In other words, the core engine is configured to operate the HP spool at a maximum speed corresponding to 85% of the rated power output of the propulsion system 10. The propulsion system 10 is also configured to produce up to 15% of the rated power output of the propulsion system via the LP spool and the inter-turbine combustor using gaseous fuel and combustion gases produced from the core engine. Such a ratio may allow for a significant reduction in the heat load imparted to downstream turbine components from the combustion section 26, which may allow for improved durability and reduced cooling flow, which may improve overall propulsion system efficiency.

In another embodiment, the propulsion system is configured to produce 80/20 work distribution between the core engine and the LP spool and inter-turbine combustor. In yet another embodiment, the propulsion system is configured to produce 75/25 work distribution between the core engine and the LP spool and inter-turbine combustor. In yet another embodiment, the propulsion system is configured to produce 60/40 work distribution between the core engine and the LP spool and the inter-turbine combustor. In still other various embodiments, the propulsion system is configured to generate between 60% and 85% of the maximum power output by combustion gases produced by the core engine, and to generate the remainder of the maximum power output using the gaseous fuel and the combustion gases produced by the core engine via the LP spool and the inter-turbine combustor.

In various embodiments, the power split is between the remaining difference of the low power output and the maximum power output. In other words, the work distribution is a limit between low power output operating conditions, above which (via the inter-turbine combustor and gaseous fuel flow) the operating conditions are high power output conditions. In certain embodiments, the maximum power output is in particular a nominal power output referring to a maximum rotational speed of the propulsion system when in normal operation. For example, the propulsion system may be operated at a rated speed or a rated power output during maximum load operation (e.g., during takeoff operation with respect to a Landing Takeoff (LTO) cycle). In certain embodiments, the limit or demarcation of the power allocation of the maximum power output (e.g., typically 60% -85%, such as 85%, or 80%, or 75%, or 60%) corresponds to cruise or descent operation of the propulsion system and the aircraft relative to the difference in the LTO cycle and the rated power output of the propulsion system. Thus, certain embodiments of the propulsion system are configured for operation from a first fuel system that provides only liquid fuel (i.e., operation without an inter-turbine combustor) to a maximum rotational speed corresponding to operation of the core engine at cruise conditions. In certain embodiments, the propulsion system is configured for maximum or rated power output from operation of both the combustion section with the first fuel system and the inter-turbine combustor with the second fuel system.

It should be understood that the range and ratio of work distribution provided herein corresponds to the particular configuration and dimensions of the core engine, the inter-turbine combustor, and the LP spool, as understood by those skilled in the art. A typical aircraft gas turbine propulsion engine is designed, sized and configured to produce 100% of the maximum power output via the combustion gases produced in the combustion section and extracted via the LP spool. Some aircraft gas turbine propulsion engines utilize an afterburner or reheat system configured to utilize a portion of liquid fuel that is typically directed to the main combustor in the combustion section and mixed with the combustion gases downstream of the main combustor to further generate thrust (i.e., an afterburner). However, such typical afterburner systems are generally not suitable for use with commercial aircraft or other aircraft that are limited by emissions output. Furthermore, such systems utilizing liquid fuels are often complex, with igniter systems and complexities associated with the lower flammability of the liquid fuels. Such systems typically produce a level of emissions, smoke or noise that may be prohibitive for use on commercial aircraft.

Referring back to fig. 2, propulsion system 10 may also include a computing system 210, computing system 210 being configured to operate propulsion system 10 such as described herein. Computing system 210 may correspond to any suitable processor-based device, including one or more computing devices, such as described above. In certain embodiments, computing system 210 is a Full Authority Digital Engine Controller (FADEC) for a gas turbine engine, or other computing module or controller configured to execute instructions for operating a gas turbine engine. For example, FIG. 2 illustrates one embodiment of suitable components that may be included within computing system 210. The computing system 210 may include a processor 212 and associated memory 214 configured to perform various computer-implemented functions.

As shown, computing system 210 may include control logic 216 stored in memory 214. The control logic 216 may include instructions that, when executed by the one or more processors 212, cause the one or more processors 212 to perform operations. In addition, computing system 210 may also include a communication interface module 230. In several embodiments, the communication interface module 230 may include associated electronic circuitry for transmitting and receiving data. Accordingly, communication interface module 230 of computing system 210 may be used to transmit data to propulsion system 10 and/or receive data from propulsion system 10. Additionally, communication interface module 230 may also be used to communicate with any other suitable component of propulsion system 10, such as described herein.

It should be appreciated that the communication interface module 230 may be any combination of suitable wired and/or wireless communication interfaces and, thus, may be communicatively coupled to one or more components of the power generation system via a wired and/or wireless connection or distribution network. The communication interface module 230 may include any suitable wired and/or wireless communication link for transferring communications and/or data, as described herein. For example, module 230 may include a SATCOM network, an ACARS network, an ARINC network, a SITA network, an AVICOM network, a VHF network, an HF network, a Wi-Fi network, a WiMAX network, a gatelink network, and the like.

A method for operating a propulsion system of an aircraft (hereinafter "method 1000") is provided. The method may be performed using an aircraft and propulsion system as described above or other suitable system. In particular embodiments, method 1000 may be performed using a computing system 210 of a propulsion system 10 or aircraft 100, such as a computer-implemented method. It should be appreciated that the computing system 210 and method 1000 provided herein may allow for improved propulsion efficiency, reduced emissions output, and overall improvements in engine and aircraft operation. Certain embodiments may provide benefits specific to propulsion systems and aircraft under limits in emissions output, noise, or thrust.

Method 1000 includes, at 1010, flowing a liquid fuel stream to a combustion section of a propulsion system. Method 1000 includes, at 1020, generating a first combustion gas corresponding to 85% or less of a rated power output of the propulsion system in the combustion section. Method 1000 includes, at 1030, modulating a rotational speed of the LP spool via modulating a flow of gaseous fuel to the inter-turbine combustor to generate a second combustion gas, such as depicted and described herein.

In various embodiments, method 1000 includes operating the core engine and the Low Pressure (LP) spool with the inter-turbine combustor at 1022 at a rated power output ratio of the core engine to the inter-turbine combustor with the LP spool, between 1.5 and 5.7, as described above. In a particular embodiment, the method 1000 includes operating the High Pressure (HP) spool at 1024 at a maximum rotational speed corresponding to between 60% and 85% of a rated power output of the propulsion system. Method 1000 includes, at 1026, flowing the gaseous fuel to the inter-turbine combustor to produce a rated power output of the propulsion system. Thus, method 1000 may operate the engine at substantially steady state operating conditions via a liquid fuel stream up to 60% to 85% of rated power output, and method 1000 may modulate the gaseous fuel stream to produce the remainder of the rated power output or a portion thereof.

In certain embodiments, method 1000 includes operating the High Pressure (HP) spool at a steady state speed at 1040 while modulating gaseous fuel flow to the inter-turbine combustor. In certain embodiments, the operation includes receiving, at 1042, a control signal corresponding to a high power mode of operation of the propulsion system. In some embodiments, receiving the control signal corresponding to the high power mode of operation includes a power rated operation or a takeoff mode of operation of the propulsion system. In other embodiments, the high power mode of operation corresponds to a climb, descent, or approach or takeoff condition relative to the LTO cycle. The method 1000 includes, at 1044, flowing the gaseous fuel to the inter-turbine combustor to produce a second combustion gas corresponding to a difference between a rated power output of the propulsion system and a power output produced by flowing the liquid fuel to the combustion section.

In another embodiment, the method 1000 includes receiving 1046 a control signal corresponding to a low power mode of operation of the propulsion system. In a particular embodiment, the low power mode of operation corresponds to cruise conditions relative to the LTO cycle. Method 1000 includes, at 1048, reducing a flow of gaseous fuel to the inter-turbine combustor to reduce a power output of the propulsion system. In a particular embodiment, method 1000 includes, at 1050, operating a High Pressure (HP) spool at a steady state rotational speed while reducing a flow of gaseous fuel to an inter-turbine combustor. In yet another particular embodiment, reducing the flow of gaseous fuel to the inter-turbine combustor to reduce the power output of the propulsion system corresponds to changing the operating mode of the propulsion system from a high power operating mode to a low power operating mode.

It should be understood that one skilled in the art will understand the elapsed time and tolerance, range, or deviation for a given speed or power output corresponding to a "steady state" operating condition. In certain embodiments, those skilled in the art will understand "steady state" in the context of an aviation propulsion system. In yet another particular embodiment, one of ordinary skill in the art will appreciate the "steady state", speed, or power output provided herein in the context of a landing takeoff cycle of an aircraft.

It should be appreciated that embodiments of the propulsion system 10, aircraft 100, and method 1000 provided herein include a combination of elements, subsystems, arrangements, and configurations that provide unexpected benefits over known elements, alone or in known arrangements and configurations. For example, it should be appreciated that having separate fuel systems and methods for control, such as via first and

second fuel systems

350 and 360, as well as method 1000 provided herein, introduces elements that may have heretofore been considered additional complications or diligence, such as discouragement of implementation in certain propulsion systems and aircraft (e.g., commercial or general aviation aircraft). However, as provided herein, the present disclosure describes systems, methods, and specific combinations or arrangements that provide unexpected benefits beyond the complexity associated with individual fuel systems.

Such benefits include allowing substantially steady state rotation speeds or operation of the HP spools while increasing and decreasing the power output of the propulsion system. Such benefits may allow for operation of one or more propulsion systems of an aircraft to generate electrical power for aircraft subsystems during idle operating conditions, runway taxi or gate-side operations, or other situations where known aircraft propulsion systems may be inoperable due to higher fuel consumption than when using an Auxiliary Power Unit (APU) to generate electrical power for an aircraft or other propulsion system. Accordingly, embodiments of the propulsion system and engine provided herein may eliminate the need or desire for an APU in an aircraft, for example, to reduce aircraft weight and increase aircraft efficiency.

Such benefits may also include allowing the core engine to have a smaller size and less fuel consumption to produce the rated power output of known propulsion systems having relatively larger core engines. The embodiments provided herein allow the core engine of the propulsion system to perform more typical APU operations and are distinct from the operations typically performed for aircraft propulsion systems. Additionally, the embodiments provided herein allow for improved emissions output over known propulsion systems, such as via reduced core engine size and improved emissions output from gaseous fuels to produce rated power output at particular engine operating conditions. Still further, by providing the second fuel system 360 and method 1000 for operating under specific operating conditions, the problems associated with gaseous fuels are alleviated as compared to using gaseous fuels for substantially all operating conditions.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Further aspects of the invention are provided by the subject matter of the following clauses:

1. an aircraft propulsion system, the propulsion system comprising: a Low Pressure (LP) spool comprising a fan section, an LP compressor, and an LP turbine; a core engine including a High Pressure (HP) compressor, a combustion section, and a HP turbine, wherein the HP compressor and the HP turbine together form a rotatable HP spool; a frame positioned in a serial flow arrangement between the HP turbine and the LP turbine, wherein the frame includes an inter-turbine combustor including struts that form exit openings into a core flow path of the propulsion system; a first fuel system comprising a first fuel conduit in fluid communication with a fuel nozzle at the combustion section, wherein the first fuel system is configured to flow a liquid fuel to the combustion section to produce a first combustion gas; a second fuel system comprising a second fuel conduit in fluid communication with the core flow path via the outlet opening at the inter-turbine combustor, wherein the second fuel system is configured to flow a gaseous fuel to the core flow path to produce a second combustion gas; wherein the LP compressor, the HP compressor, the combustion section, the HP turbine, the inter-turbine combustor, and the LP turbine are in a series flow arrangement; and wherein said propulsion system comprises a rated power output ratio of said core engine to said inter-turbine combustor with said LP spool between 1.5 and 5.7.

2. The propulsion system of any one or more of the clauses herein, wherein the core engine is configured to operate the HP spool at a maximum rotational speed corresponding to between 60% and 85% of the rated power output of the propulsion system.

3. The propulsion system of any one or more of the clauses herein, comprising: a computing system comprising a processor and a memory, wherein the memory is configured to store instructions that, when executed by the processor, cause the propulsion system to perform operations comprising: flowing a liquid fuel to the combustion section and then generating a first combustion gas in the combustion section corresponding to 85% or less of the rated power output of the propulsion system.

4. The propulsion system of any one or more of the clauses herein, the operations comprising: modulating a rotational speed of the LP spool via modulating the flow of gaseous fuel to the inter-turbine combustor to generate the second combustion gas.

5. The propulsion system of any one or more of the clauses herein, the operations comprising: operating the HP spool at a steady state rotational speed while modulating the flow of gaseous fuel to the inter-turbine combustor.

6. The propulsion system of any one or more of the clauses herein, the operations comprising: modulating gaseous fuel flow through the inter-turbine combustor to vary an output power of the propulsion system.

7. The propulsion system of any one or more of the clauses herein, the operations comprising: maintaining a steady state rotational speed of the HP spool while modulating the flow of gaseous fuel.

8. The propulsion system of any one or more of the clauses herein, the operations comprising: receiving a control signal corresponding to a high power mode of operation of the propulsion system; gaseous fuel is then flowed to the inter-turbine combustor to generate the second combustion gases corresponding to a difference between the rated power output of the propulsion system and a power output generated by flowing liquid fuel to the combustion section.

9. The propulsion system of any one or more of the clauses herein, the operations comprising: receiving a control signal corresponding to a low power mode of operation of the propulsion system; the flow of gaseous fuel to the inter-turbine combustor is then reduced, thereby reducing the power output of the propulsion system.

10. The propulsion system of any one or more of the clauses herein, the operations comprising: operating the HP spool at a steady state rotational speed while reducing the flow of gaseous fuel to the inter-turbine combustor.

11. The propulsion system of any one or more of the clauses herein, wherein the fan section is configured as a ductless open rotor.

12. The propulsion system according to any one or more of the clauses herein, comprising: a motor operably coupled to the HP spool.

13. A computing system for an aircraft propulsion system, the computing system comprising one or more processors and one or more memories, wherein the memories are configured to store instructions that, when executed by the processors, cause the propulsion system to perform operations comprising: flowing a liquid fuel stream to a combustion section of the propulsion system; generating a first combustion gas corresponding to 85% or less of a rated power output of the propulsion system in the combustion section; and modulating a rotational speed of the LP spool via modulating a flow of gaseous fuel to an inter-turbine combustor to produce a second combustion gas.

14. The computing system of any one or more of the clauses herein, the operations comprising: operating a High Pressure (HP) spool at a steady state rotational speed while modulating the flow of gaseous fuel to the inter-turbine combustor.

15. The computing system of any one or more of the clauses herein, the operations comprising: receiving a control signal corresponding to a high power mode of operation of the propulsion system; and flowing a gaseous fuel to the inter-turbine combustor to produce the second combustion gas corresponding to a difference between the rated power output of the propulsion system and a power output produced by flowing a liquid fuel to the combustion section.

16. The computing system of any one or more of the clauses herein, the operations comprising: receiving a control signal corresponding to a low power mode of operation of the propulsion system; and reducing the flow of gaseous fuel to the inter-turbine combustor, thereby reducing the power output of the propulsion system.

17. The computing system of any one or more of the clauses herein, the operations comprising: operating a High Pressure (HP) spool at a steady state rotational speed while reducing the flow of gaseous fuel to the inter-turbine combustor.

18. The computing system of any one or more of the clauses herein, the operations comprising: operating the core engine and the Low Pressure (LP) spool with the inter-turbine combustor at a rated power output ratio of the core engine to the inter-turbine combustor with the LP spool between 1.5 and 5.7.

19. The computing system of any one or more of the clauses herein, the operations comprising: operating a High Pressure (HP) spool at a maximum rotational speed corresponding to between 60% and 85% of the rated power output of the propulsion system.

20. The computing system of any one or more of the clauses herein, the operations comprising: flowing a gaseous fuel to the inter-turbine combustor to produce the rated power output of the propulsion system.

21. A propulsion system according to any one or more of the clauses herein, including the computing system of any one or more of the clauses herein.

22. The computing system of any one or more of the clauses herein configured to operate the propulsion system of any one or more of the clauses herein.

23. An aircraft comprising the propulsion system of any one or more of the clauses herein.

24. An aircraft comprising the computing system of any one or more of the clauses herein.

Claims

1. An aircraft propulsion system, characterized in that the propulsion system comprises:

a Low Pressure (LP) spool comprising a fan section, an LP compressor, and an LP turbine;

a core engine including a High Pressure (HP) compressor, a combustion section, and an HP turbine, wherein the HP compressor and the HP turbine together form a rotatable HP spool;

a frame positioned in a serial flow arrangement between the HP turbine and the LP turbine, wherein the frame includes an inter-turbine combustor including struts that form outlet openings into a core flow path of the propulsion system;

a first fuel system comprising a first conduit in fluid communication with a fuel nozzle at the combustion section, wherein the first fuel system is configured to flow a liquid fuel to the combustion section to produce a first combustion gas;

a second fuel system comprising a second conduit in fluid communication with the core flow path via the outlet opening at the inter-turbine combustor, wherein the second fuel system is configured to flow a gaseous fuel to the core flow path to produce a second combustion gas;

wherein the LP compressor, the HP compressor, the combustion section, the HP turbine, the inter-turbine combustor, and the LP turbine are in a series flow arrangement; and is

Wherein the propulsion system includes a rated power output ratio of the core engine to the inter-turbine combustor with the LP spool between 1.5 and 5.7.

2. The propulsion system of claim 1, wherein the core engine is configured to operate the HP spool at a maximum rotational speed corresponding to between 60% and 85% of the rated power output of the propulsion system.

3. A propulsion system as in claim 1 wherein the propulsion system comprises:

a computing system comprising a processor and a memory, wherein the memory is configured to store instructions that, when executed by the processor, cause the propulsion system to perform operations comprising:

flowing a liquid fuel to the combustion section and then generating a first combustion gas in the combustion section corresponding to 85% or less of the rated power output of the propulsion system.

4. A propulsion system as in claim 3 wherein the operations comprise:

modulating a rotational speed of the LP spool via modulating a flow of gaseous fuel to the inter-turbine combustor to generate the second combustion gas.

5. A propulsion system as in claim 4 wherein the operations comprise:

operating the HP spool at a steady state rotational speed while modulating the flow of gaseous fuel to the inter-turbine combustor.

6. A propulsion system as in claim 3 wherein the operations comprise:

modulating a flow of gaseous fuel through the inter-turbine combustor to vary a power output of the propulsion system.

7. A propulsion system as in claim 6 wherein the operations comprise:

maintaining a steady state rotational speed of the HP spool while modulating the flow of gaseous fuel.

8. A propulsion system as in claim 3 wherein the operations comprise:

receiving a control signal corresponding to a high power mode of operation of the propulsion system; then the

Flowing gaseous fuel to the inter-turbine combustor to generate the second combustion gas corresponding to a difference between the rated power output of the propulsion system and a power output generated by flowing liquid fuel to the combustion section.

9. A propulsion system as in claim 3 wherein the operations comprise:

receiving a control signal corresponding to a low power mode of operation of the propulsion system; then the

Reducing the flow of gaseous fuel to the inter-turbine combustor, thereby reducing the power output of the propulsion system.

10. A propulsion system as in claim 9 wherein the operations comprise:

operating the HP spool at a steady state rotational speed while reducing the flow of gaseous fuel to the inter-turbine combustor.